CONTENTS



Power, thermal, and environmental design

Presented by:

Mike Belcher

Joe Bidwell

Chris Campbell

Miriam Scheidt

Chris Zuchowski

AOE 4065

Dr. C.D. Hall

November 16, 2001

Contents

List of Figures iii

List of Tables iii

List of Symbols iv

Greek v

List of Abbreviations vi

Chapter 1: Introduction 1

1.1 Power Systems17 1

1.1.1 Solar Panels 1

1.1.2 Energy Storage 2

1.1.3 Tether-Specific Application 4

1.3 Thermal System13, 29 4

1.3.1 Sources of Thermal Radiation 4

1.3.2 Thermal Components 4

1.4 Space Environment 5

1.4.1 Effects of the Atmosphere 6

1.4.2 Effects of Thermal Radiation 6

1.4.3 Effects of Ionization 6

1.4.4 Effects of Plasma Charging 8

1.4.5 Effects of Debris 8

1.5 Summary and Overview 9

Chapter 2: System Modeling 10

2.1 Power System 10

2.1.1 Power Generation 10

2.1.2 Energy Storage 12

2.2 Thermal 15

2.3 Orbital Debris29 18

2.4 Solar Particle Events 21

2.5 Single-Event Phenomena 21

2.6 Earth’s Magnetic Field 22

2.7 Interaction Matrix 23

2.8 Intersystem Interactions 24

2.9 Summary and Overview 29

Chapter 3: Subsystem Examples 30

3.1 Power Generation Components 30

3.2 Power Generation Example Calculations 32

3.3 Energy Storage Components 32

3.4 Thermal Subsystem Components 36

3.5 Thermal Subsystem Calculations 38

3.6 Environment Components 41

3.6.1 Single Particle Events 41

3.6.2 Electronic Devices 42

3.6.3 Debris and Thermal Protection 43

3.6.4 Radiation Detection 43

3.7 Environment Design Calculations 44

3.8 Summary and Overview 45

Chapter 4: Conclusions 46

4.1 Summary and Conclusions 46

4.2 Recommendations 48

References 50

List of Figures

Figure 1: Cycle Life vs. Depth of Discharge

Figure 2: Spatial Density Values

Figure 3: Time Evolution of a Solar Particle Event Observed on Earth

Figure 4: Subsystem Interactions

Figure 5: Schematic of Spacecraft Used for Example Calculations

Figure 6: Radiation Dose Rates as a Function of Altitude29

List of Tables

Table 1: Collision Probability per Year29

Table 2: Nickel Cadmium Cylindrical Rechargeable Battery32

Table 3: Nickel Hydride Cylindrical Rechargeable Battery32

Table 4: Lithium-Ion Cylindrical Rechargeable Battery32

Table 5: Conduction Couplings Example

Table 6: Radiation Couplings Example

List of Symbols

A Surface area

Ac Collision cross section

Asa Area of solar array

a, b Surface dimensions

B Local magnetic field

Bo Earth magnetic field at equator

c Speed of light, depth

C Capacity

Cd Drag coefficient

cg Center of mass

cp Specific heat

cpa Center of pressure

D Residual dipole of spacecraft

F View factor

Fs Solar constant

h Heat transfer coefficient

i Sun angle of incidence

I Moment of inertia

Id Inherent degradation

k Thermal conductivity

Ld Lifetime degradation factor

m Mass

n Load transmission efficiency

N Number of batteries

Pc Probability of debris impact

Pd Spacecraft power requirement during daylight

Pe Spacecraft power requirement during eclipse

Po Solar cell output per unit area

Psa Power produced by solar array

PBOL Solar array power per unit area at spacecraft beginning of life

PEOL Solar array power per unit area at spacecraft end of life

q Coefficient of reflectivity

Q Rate of heat flow

Qc Critical charge

R Orbit radius, number of errors per bit day

RE Earth radius

S Wall spacing

t Duration of mission

T Temperature

Ta Aerodynamic torque

tb Thickness of backup layer

Td Time spent in daylight

Te Time spent in eclipse

TE Gravity gradient torque

Tm Magnetic Torque

Tsp Solar pressure torque

ΔT Temperature difference

v Velocity

Vrel Relative velocity between spacecraft and debris

Xd Efficiency of electrical path directly from solar arrays to loads

Xe Efficiency of electrical path from solar arrays through batteries to individual loads

Δx Path length of heat transfer

Greek

α Absorptivity

ε Emissivity

θ Sun incidence angle, maximum angular deviation of z axis from local vertical

λ Magnetic latitude

μ Gravitational parameter

σ Stefan-Boltzmann constant

List of Abbreviations

ADCS Attitude Determination and Control System

BOL Beginning of Life

CEASE Compact Environmental Anomaly Sensor

CREME Cosmic Ray Effects on Microelectronics

DOD Depth-of-Discharge

EOL End of Life

GaAs Gallium-Arsenide

GEO Geostationary Orbit

HTPC High Temperature Phase Change

IR Infrared Radiation

ISS International Space Station

LEO Low-Earth Orbit

LET Linear Energy Transfer

Li-Ion Lithium-Ion

MLI Multi-Layer Insulation

Ni-Cd Nickel-Cadmium

NiH2 Nickel-Hydrogen

Ni-MH Nickel-Metal Hydride

NSSDC National Space Science Data Center

RTG Radioisotopic Thermoelectric Generator

SEP Single Event Phenomena

SEU Single Event Upset

SPE Single Particle Event

SSN Space Surveillance Network

Chapter 1: Introduction

The power system of a spacecraft is a crucial element of its design. Inefficiencies in the design of the power system can add unnecessary mass, fill valuable space, and cause less than optimal performance of the spacecraft as a whole. Of additional importance to a spacecraft’s design are the thermal requirements and the expected space environment concerns. It is essential that the thermal system of the spacecraft maintain component temperatures within their limits. Both active and passive temperature control measures exist, and an optimal design will minimize the mass and power required by the temperature control components. Protection from the space environment is also necessary to ensure spacecraft components functionality for the duration of the mission. Potential hazards to a spacecraft include solar radiation, solar events, orbital debris and micrometeoroids. However, knowing of these possible factors, a spacecraft’s design can safeguard against their effects. In this chapter, an introduction is given to power generation, energy storage, thermal systems, and the properties of the space environment.

1.1 Power Systems17

1.1.1 Solar Panels

Spacecraft power generation usually comes from a renewable source of energy, such as the sun. Most satellites and even deep space probes use solar arrays that harness the sun’s energy, converting it to power they require for internal functions. These solar arrays are usually comprised of layered materials such as gallium and arsenide, each of which absorbs different wavelengths of the sun’s radiation to provide power. Each panel typically consists of many panels, the constituents of which are even smaller solar cells. Individually, the solar cells only generate a few watts, but when large quantities are placed in arrays, kilowatts of power can be produced, enough to maintain many onboard processes. For maximum efficiency, however, the panels must be perpendicular to the direction the sun’s radiation is traveling, and as such the solar panels are commonly movable.

Two of the most common types of solar cells in production today are the crystalline silicon and gallium arsenide cells. Though gallium-arsenide cells are approximately twice as efficient as silicon cells, they must also be manufactured especially for photovoltaic use, and are therefore much more expensive.

1.1.2 Energy Storage

The energy needed to run a satellite must be included in the bus or captured from the sun with the use of solar cells. In either case, a system for storing this energy is needed. If solar cells are not used, then all the energy required by the satellite for its lifetime must be stored on the satellite at launch. If solar cells are used for harnessing the energy emitted by the sun, a storage system is needed for when the earth eclipses the satellite. The satellite would then store extra energy when the sun is visible and use that energy when sunlight can be collected.

The three major ways to store energy on a satellite are by using batteries, flywheel energy storage, and thermal energy storage. The objectives and constraints of the satellite determine which of these methods is used. They can be used separately or in conjunction with one another.

The most common energy storage system is batteries. Batteries use a chemical reaction to produce a flow of electrons between negative and positive electrodes. When the device being powered is connected to these electrodes the reaction begins and disconnecting the electrodes ends the reaction. Primary batteries are storage devices in which the reaction happens only once. These batteries are good for single-use operations. However, for satellites with longer lifetimes and requirement to operate in eclipse, secondary batteries are necessary. These batteries are similar to primary batteries, though passing electricity back through the batteries can recharge them. This process allows these devices to be used repeatedly. Batteries are named for their chemical composition. The most widely space-used batteries are Silver-Zinc, Nickel-Cadmium, Nickel-Hydrogen, and Lithium. All of these batteries work in the same general way, but use a different chemical reaction.

New practice allows flywheels to be used as energy storage systems. These storage systems store energy collected in the form of kinetic energy. The excess energy collected is used to spin a flywheel on a magnetic bearing to a high angular velocity. The kinetic energy of the spinning wheel is then converted back to electricity when it is needed. A flywheel energy storage system can also be used for attitude control. Thermal energy storage systems use high-temperature phase-change (HTPC) materials to store excess energy collected. The energy is used to melt the HTPC materials. The material stays in the liquid phase until the energy is needed. When needed, it melts and the heat energy is converted to electricity by heat engines.

1.1.3 Tether-Specific Application

Aboard a tethered spacecraft, solar arrays will generally be used to provide power to batteries, or some sort of energy storage devices, so that they may provide current to the tether when needed. Through the clever exploitation of electrodynamics, this current will generate a force on the tether as it orbits through the Earth’s magnetic field. For this application, we will need to have a power source onboard the space tether, and solar power is currently the most efficient form for use on a spacecraft.

1.3 Thermal System13, 29

1.3.1 Sources of Thermal Radiation

The purpose of the thermal subsystem is to keep the spacecraft’s vital instruments within operational temperature limits. In space, a satellite is subject to heat input from several sources, including the sun, infrared radiation from the Earth, and power dissipation of the onboard components. Additional consideration must be given to an albedo factor, which is a percentage of solar radiation reflected by the Earth’s atmosphere. The spacecraft also loses heat to space. Thus, the thermal subsystem must take into account many factors to keep the spacecraft within temperature specifications.

1.3.2 Thermal Components

The thermal subsystem can consist of several components of varying complexity. Among the simplest are thermal finishes and Multi-Layered Insulation (MLI). Thermal finishes include paints, silverized plastics and coatings such as anodizing. These finishes improve thermal performance by altering the radiative properties of the material they cover. MLI, also known as thermal blanketing, insulates the spacecraft by minimizing radiation heat transfer from a component to its surroundings. Forms of more active thermal control systems are thermostats, heaters and radiators. A thermostat works in conjunction with a heater by monitoring the temperature of a specific component and then turning the heater attached to it on or off according to preset operational constraints. A radiator receives heat from components, usually via heat pipes, and then emits the heat to space. It is usually part of the external spacecraft structure and is usually coated with a highly emissive material. A heat pipe is a sealed tube containing a phase-changing liquid that allows it to carry heat from a source to a sink. Other, less common, thermal control devices include louvers and cryogenic systems. Louvers resemble Venetian blinds that can open and close and thus regulate the rate at which a radiator dissipates heat to space. Cryogenic systems typically use cold liquids, usually liquid nitrogen, to keep sensitive components such as IR detectors cool. Analyzing, choosing, and successfully implementing some of the subsystem components listed here are necessary for maintenance of thermal control of a design.

1.4 Space Environment

The natural space environment contains many concerns that need to be considered in the design of a spacecraft. Many properties of space can be detrimental to components of a satellite. Most of these characteristics are difficult to predict because they vary with the solar events that change with time of day, season, and the rotation of the sun.29 The environmental effects that must be considered include:

• Earth’s atmosphere

• Thermal radiation

• Ionization

• Plasma charging

• Meteorites and orbital debris

1.4.1 Effects of the Atmosphere

In orbits close to the Earth, the effects of the atmosphere must be considered. The atmosphere creates drag on the spacecraft, which degrades the orbit.8 Particles in the atmosphere also erode the surfaces of the satellite through collisions at speeds of about 8 km/s.2 This phenomenon can affect the thermal and electrical properties and degrade the structure itself. However, the atmosphere can also offer some protection to the spacecraft against thermal and ionizing radiation in altitudes up to 200 km.20

1.4.2 Effects of Thermal Radiation

Satellites are exposed to thermal energy from three sources: (1) incoming solar radiation, (2) reflected solar energy, and (3) outgoing longwave radiation from the Earth and its atmosphere. Repeated changes in temperature will lead to fatigue in spacecraft components.8

1.4.3 Effects of Ionization

Ionization is a problem in most regions of the space environment. In the atmosphere, free atomic oxygen, when combined with ultraviolet radiation can manipulate the mechanical, optical, and thermal properties of surfaces.8 In the upper atmosphere, there is a small region that consists of positive ions and free electrons. Traveling through this region causes disruptions in many electronic components.8,21 The near-vacuum conditions of space cause organic materials to outgas, meaning that they create spurious molecules which contaminate other surfaces of the satellite.29 Ionizing radiation comes from three sources: (1) trapped radiation belt particles (the Van Allen belts), (2) cosmic rays, and (3) solar flares.8 This radiation decreases power production, causes failure of electronic components such as electric circuits and solar panels, and increases background noise in sensors.

Solar particle events (SPE) tend to occur in association with solar flares, and last between several hours and several days.15 An SPE consists of a rapid increase in the flux of energetic particles. Damage to solar arrays can be reduced through choice of cell type and cover glass thickness.15 Effects from SPEs can also be minimized by choosing low inclination orbits where only particles with sufficiently high energies are able to penetrate through the Earth’s magnetic field.29

Shielding from radiation is important in two ways. The most important consideration in manned missions is the amount of radiation to which humans can safely be exposed.20 Furthermore, modern electronics are more sensitive to radiation and the single event phenomena that may occur, and therefore need more shielding.

Due to the miniscule dimensions of some modern day electronic devices, a single energetic particle may be able to cause a small mutation in current capable of flipping a memory bit. Sometimes the effect is only temporary, but in its worst case, the event may cause “latchup.” Latchup is a situation in which a bit is flipped causing a device to become isolated to new input. Typical circuits have a latchup threshold on the order of 108 Rad/s.14 “Upset” is a slightly different case in which the operation of a device is altered from its design specifications. Upset tends to occur from a threshold of 107 Rad/s up to 1012 Rad/s.14 Generally, the more likely a device is to absorb radiation, the more likely it is to suffer from a latchup or upset.14

1.4.4 Effects of Plasma Charging

The plasma environment can be another hazard to spacecraft. One of the major effects of the plasma environments is the slow accumulation of charge on the surfaces of a spacecraft. The electrostatic fields resulting from the surface charging can have detrimental effects on the spacecraft. Surface arc discharges can generate electromagnetic interference and induce currents in electrical systems.29 Prolonged charging can cause changes in surface, thermal, and optical properties and can also shift the spacecraft electrical ground.29 Careful selection of materials and conductive coatings, and the use of special cabling or grounding can be used to inhibit spacecraft charging. In addition, low altitude, low inclination orbits generally have a lower potential for spacecraft charging.28

1.4.5 Effects of Debris

The last consideration in the space environment examined here is the impact of micrometeorites and orbital debris. Micrometeorites are remnants of comets, whereas orbital debris is man-made and consists of parts of obsolete or damaged spacecraft. All of this debris poses a threat of collision with satellites in orbit around the Earth. It is vital, therefore, to consider shielding of critical components of the spacecraft. Higher orbits may be safer as most debris is found within 2000 km of the Earth’s surface, with high concentrations found at 800 km, 1000 km, and 1500 km.28

1.5 Summary and Overview

This chapter gives an overview of the different elements to be considered in designing a power, thermal, and space environment system. The remaining chapters will discuss modeling and analysis techniques and give hypothetical and historical examples of previous designs relevant to these systems.

Chapter 2: System Modeling

To analyze a subsystem, various equations must be used to model performance characteristics of the system. Analysis includes utilizing equations and functions that model potential spacecraft configurations. In addition, interactions between subsystems must be considered to determine mutually exclusive design choices.

2.1 Power System

Power generation and energy storage are the two main components of the power system. Consideration towards both of these systems should be done concurrently during the design phase. The advantages and disadvantages of both the power generation and energy storage methods must both be considered together in order to arrive at an optimal design.

2.1.1 Power Generation

Without generation of sufficient power for a spacecraft, critical operations may not be performed, possibly resulting in a mission failure. For this reason, the spacecraft bus must provide adequate power to its components at all times. Many different forms of power generation exist, including solar photovoltaic, solar thermal dynamic, radioisotope, nuclear, and fuel cells. We must first establish the power requirements and when during the spacecraft’s orbit power is required to determine which of these power sources is optimal. Another factor that must be considered in power source selection is the essential spacecraft lifetime. Many of the strengths and weaknesses of the varying power sources can be assessed by referencing Table 11-33 in Ref. 29.

Many of the satellites in orbit today require some form of regenerative power supply. Currently the most efficient way of providing a nearly limitless supply of power is using solar cells. Solar cells are the most commonly used form of power generation among orbital spacecraft today because of their low mass and cost. The amount of power to be produced by the solar array, Psa, is used to determine the size of the array and is calculated using the following equation:

(2.1-1)

where Pe and Pd are the spacecraft power requirements during eclipse and daylight, respectively, and Te and Td are the lengths of these periods per orbit. The parameters Xe and Xd are the efficiencies of the paths from the solar arrays through the batteries to the individual loads and the path directly from the arrays to the loads, respectively.29 Using the value Psa helps determine which solar panel material will be used. Silicon, thin sheet amorphous silicon, gallium arsenide, indium phosphide, and multijunction GaInP/GaAs cells are among the current options available for solar photovoltaic materials. Table 11-35 in Ref. 29. lists a performance comparison between the different photovoltaic solar cells, including their energy-conversion efficiency. Using this efficiency, we can calculate the ideal solar cell output per unit area, Po by multiplying the efficiency by the potential solar radiation, 1367 W/m2. Next, we have to ascertain the realistic power production capability of the solar array. Factors degrading the solar array efficiency further are the design and assembly of the cells, the operating temperature of the array, and the shadowing of the cells in space. These three factors are combined in the inherent degradation, Id, with degradation values typically ranging from 0.49 to 0.88. Given Po, Id, and the sun incidence angle (, we can calculate the array’s power per unit area at its beginning of life, PBOL, using the following equation:

PBOL= PoIdcos ( (2.1-2)

The sun incidence angle is the angle between the vector normal to the surface of the array and the sun line. In calculations, typically the worst-case situation is used. The life of solar arrays is limited by orbital radiation. Usually this radiation damage is manifested in a lifetime degradation factor, Ld, which can be calculated from the following equation:

Ld=(1-degradation/year)satellite life (2.1-3)

The solar array degradation per year can be found in Table 11-35 in Ref. 29. With the degradation factor, we can determine the array’s performance per unit area at the end of its life, PEOL by using this formula:

PEOL= PBOLLd (2.1-4)

Finally, the solar array area, Asa, required to support the spacecraft’s power requirement, Psa, is given by:

Asa= Psa/PEOL (2.1-5)

Using this method, we can easily and accurately determine how large we need to make a particular solar array to power the spacecraft even after years of degradation effects.

2.1.2 Energy Storage

Excess energy collected must be stored so that it can be used during the eclipse period. The eclipse period is when the satellite passes behind the earth relative to the sun. During this time, the solar cells are not capturing energy from the sun, and the satellite must operate on stored energy. The energy storage system must meet all constraints and objectives of the design. A constraint of a satellite design is its lifetime requirement. The main criteria for the design for energy storage are to maximize performance, minimize mass, and to minimize cost.

Batteries are most commonly used for energy storage in satellites. The most widely used batteries in satellites are rechargeable nickel-cadmium (Ni-Cd) or nickel-hydrogen batteries (Ni-MH). These batteries are energy cells that use a chemical reaction to create electrical energy. This type of energy storage has proven to be reliable in space applications. This attribute is important because maintenance missions in space are expensive. Nickel-hydrogen (NiH2) and Ni-Cd batteries operate optimally when kept between the temperatures of 5°C and 20°C. The lifetime of batteries is greatly influenced by the depth-of-discharge (DOD). The DOD is the amount that the battery is discharged every charge-discharge cycle. Over time, the battery remembers the DOD and will only charge the amount it is discharged. In most applications, batteries are routinely discharged completely to offset this problem. Lifetime is also influenced by temperature, rate of charge, rate of discharge, and degree of overcharging. However, these elements are not as well defined as DOD. A graph of the relationship between depth-of discharge and lifetime in terms of charge-discharge cycles for these two types of batteries is shown in Figure 1.29

[pic]

Figure 1: Cycle Life vs. Depth of Discharge29

Battery performance is rated by the amount of energy that can be stored in the cell. This measure is called the battery’s capacity. Capacity is measured in ampere-hours or watt-hours. The equation for battery capacity is:

[pic] (W-hr) (2.1-6)

In equation (2.1-6), Pe is the average eclipse load, Te is the eclipse duration, n is the load transmission efficiency, N is the number of batteries, and DOD is the depth-of-discharge. The eclipse load is defined as the power needed by the satellite during eclipse. The orbit of the satellite determines the time that the satellite is eclipsed by the earth. From Figure 1, the typical DOD for NiH2 is approximately 40-60%, and for Ni-Cd, 10-20%. In most satellite designs, there are multiple batteries in case one fails. The efficiency of the load transmission is a measure of how much power is lost when transmitted to and from the battery.

Battery mass is normally more than that of most of the other satellite components. In order to minimize mass and maximize performance, a battery must have the highest capacity per mass. This measure is called the specific energy density. The specific energy density is the battery capacity divided by the mass of the battery. Therefore, a battery with low mass and high capacity is better. Lower mass also reduces the cost of the design.29

Another energy storage option is storing energy on flywheels. Excess power produced is used to run a motor that spins a flywheel on magnetic bearings. The power is stored as kinetic energy. During eclipse, the energy in the flywheel is converted to power by running the motor backwards as a generator. This method of energy storage interacts with attitude determination and control systems (ADCS) because the spinning of the flywheel creates a moment on the satellite.23

2.2 Thermal

The thermal subsystem is responsible for keeping spacecraft components within allowable temperature limits. The allowable temperature margins of important components are required to perform the thermal analysis. In addition, information about the environmental conditions, such as solar flux, albedo, and earth IR is required. Other pertinent information includes the cost goals and the weight allocated to the thermal subsystem.

The basic laws of heat transfer govern the majority of calculations relevant to the thermal subsystem. Heat is transferred via convection, conduction and radiation.

Convection is heat transfer that occurs between a solid object and the fluid medium around it, and is governed by:

[pic] (2.1-7)

where Q is the rate of heat flow (W), h is the heat transfer coefficient (W m-2K-1), A is the surface area (m2), and ΔT is the temperature difference between the object and the fluid. Since space is a near-vacuum environment, convection does not play a large role in heat transfer of spacecraft. However, convection is important when considering atmospheric heating during launch after the payload fairing has been ejected.

Conduction describes the heat transfer that occurs between solid objects. The equation for conduction is:

[pic] (2.1-8)

where Q is the heat transfer rate (W), k is the thermal conductivity of the material

(W m-1K-1), A is the cross sectional area normal to the direction of heat transfer (m2), Δx is the path length of heat transfer (m), and ΔT is the temperature difference. Figure 11-15 in Ref. 29 illustrates how to calculate conduction for different shapes. Conduction is the dominant method of heat transfer within the spacecraft and can be used to estimate the heat transfer between objects such as the structure and computer components.

Radiation is the third basic mode of heat transfer and is the method spacecraft use to dissipate heat to space. Radiation also influences the heat exchange between internal spacecraft components. Radiation heat transfer is governed by:

[pic] (2.1-9)

where Q is the heat transfer rate (W), ε is the emissivity, the ratio of the energy emitted by a substance to the energy emitted by a blackbody at the same temperature, A is the surface area of radiating surface, σ is the Stefan-Boltzmann constant (5.67051(10-8 W/m2K4), and T2 and T1 are the temperatures of the two objects of interest.

The surface properties of a material affect how it dissipates heat. The radiative properties of a material are described by emissivity, defined above, and absorbtivity (α). Absorbtivity is defined as a percentage of incoming radiation that a material absorbs. Figure 11-44 of Ref. 29 gives values of α and ε for commonly used spacecraft materials. Another important principle in radiation calculations is the view factor (F). View factor characterizes the effects of geometry and surface orientation on the radiative heat exchange between surfaces. References 29 and 13 give more detailed information about how to calculate view factors, as well as other important information regarding the radiation heat exchange on a spacecraft.

Once the relevant heat transfer rates (Q’s) from convection, conduction and radiation are calculated for the spacecraft they can be used in the rest of the thermal analysis. The governing equation for spacecraft thermal analysis is:

[pic] (2.1-10)

which is an equilibrium energy equation. Information about the thermal environment of the satellite, as well as the thermal properties of the materials, can be used in conjunction with this equation to give steady state temperatures of components. Tables 11-47 and 11-48 in Ref. 29 give information about the thermal conductivity (k) and specific heat capacity (cp) of commonly used spacecraft materials.

Once temperatures for each of the important spacecraft components are determined, the thermal systems engineer makes changes to the thermal subsystem to ensure that components stay within specified temperature margins. Any number of thermal control devices can be used, such as radiators, thermostats and heaters, or by changing the surface coatings on the components of interest. Thus, the thermal subsystem interacts strongly with other subsystems in that changes in one will influence the other. These subsystem interactions are explained in further detail in Section 2.3.

2.3 Orbital Debris29

Orbital debris is defined as any non-operational object in space. Over 8,500 of these objects are tracked and cataloged by the Space Surveillance Network (SSN). This network consists of radar and optical facilities worldwide. The facilities are limited, however, in that the smallest objects they can sense in LEO have diameters of 10 cm and in GEO, the smallest diameter trackable is 1 m. A significant problem occurs because the largest fragment a satellite can be shielded against is about 1 cm in diameter, and small debris cannot be tracked by the SSN but can cause detrimental damage to the spacecraft. Spatial density plots, shown in Figure 2, can be used to estimate the amount of debris in space. Spatial density is the number of objects per unit volume of space. Models predicting the amount of debris are compared to special radar measurements for detectable objects and impacts caused by small debris on returned satellite surfaces.

[pic]

Figure 2: Spatial Density Values29

The impact of debris on a satellite also depends on the velocity between the two objects. In LEO, the debris’ average relative velocity to the satellite is approximately

9–10 km/s, with maximum values reaching 14 km/s in eccentric and retrograde orbits. In GEO, satellites have a lower orbital velocity; therefore, relative velocities of debris are usually 100 to 500 m/s.

The probability of debris impacting a spacecraft is estimated by the kinetic theory of gases:

[pic] (2.1-11)

where SPD is the spatial density of debris, AC is the collision cross-section, t is the duration of the mission, and Vrel is the relative velocity between the spacecraft and the debris. This equation gives an approximation of the probability of collision within one year. The order of magnitudes for this equation at a variety of altitudes and satellite sizes are given in Table 1. As seen in this table, the trackable population of debris presents a manageable hazard to even large satellites, but the probability of collision increases with the inclusion of uncataloged debris.

Table 1: Collision Probability per Year29

[pic]

Since all spacecraft will encounter debris during their lifetime, they need to be protected by either passive and/or active means. Passive protection includes shielding and redundancy, while active means avoiding a collision. A shield is designed to break apart an impacting object and to allow the smaller debris to hit backup layers over a larger area, which causes less damage. The thickness of the backup layer, tb, is determined by:

[pic] (2.1-12)

where C is an empirically derived constant of value 41.5 ±14.0 (cm3g-1km-1s), m is the projectile mass in g, v is the projectile velocity in km/s, and S is the wall spacing in cm.

2.4 Solar Particle Events

Solar particle events (SPE) are associated with solar flares and consist of rapid increases in the flux of energetic particles that can last a few hours to several days. Usually, only a few SPEs occur every year. The typical progress over time of a solar particle event is shown in Figure 3. These type of graphs depend on three factors:

1) the solar flare causing the event; 2) the length of time the energetic particles need to diffuse within the solar corona; and 3) the propagation of the particles in the interplanetary medium.

[pic]

Figure 3: Time Evolution of a Solar Particle Event Observed on Earth29

2.5 Single-Event Phenomena

Single Event Phenomena (SEP) are an important consideration to designers of electronic components on spacecraft. A rough understanding of the frequency of SEP must be reached prior to determining corrective measures. The frequency of errors to be expected is estimated using the following relation:

[pic] (2.1-13)

where R is the number of errors per bit day, a and b are device surface dimensions in (m, c is device depth in (m, and Qc is the critical charge in pC.29

Numerous methods exist for reducing problems caused by SEP. Options include the use of: redundant logic for critical functions, watchdog timers, and hard latches. Special care is given to digital circuit design and results in immunity to analog circuit spikes. The periodic refreshing of critical memories is also good practice for sound computer systems.29

2.6 Earth’s Magnetic Field

When using magnetic torquers and/or magnetometers, accurate information for the strength and direction of the Earth’s magnetic field is necessary. Magnetometers determine a spacecraft’s location and orientation by comparing magnetic field readings with previously collected data. The control laws used by magnetic torquers similarly depend on magnetic field strength information. The strength of the magnetic field can be calculated using the following relation:

[pic] (2.1-14)

where B is local magnetic field intensity, ( is the magnetic latitude, R is the radial distance measured in Earth radii (RE), and B0 is the magnetic field at the equator at the Earth’s surface (B0 = 0.30 gauss).29

2.7 Interaction Matrix

Power generation is intrinsically associated with power storage aboard any spacecraft, as the storage devices need to be sized appropriately to retain the necessary power. Stored energy is also required during periods when external power cannot be harnessed. Power generation is also moderately related to the spacecraft thermal conditions, with electrical efficiencies inversely proportional to the temperature of the craft. Radiation encountered by the spacecraft has very little effect on the ability of the power generation system. Debris can, however, present a problem for different forms of power generation, as in the case of solar arrays, and for this reason it is somewhat related to power generation.

The ADCS system relies on the energy storage system to give it power to run its components. The communications system uses the energy storage system to power its components. The storage system might interact with the communication system if a parameter in the storage system needed to be changed by command from the ground. The structural system has to support the mass of the energy storage system and the energy storage system is housed by the structures system. The mission of the satellite defines the mass, size, cost, and lifetime of the energy storage system. The components that accomplish the mission operations of the satellite may be powered by the power sub-system. Also, the orbit of the satellite determines the time of eclipse and therefore the capacity of the storage system. The propulsion system may need to be powered by the power storage system during eclipse.

The thermal subsystem interacts with many of the other subsystems. The power system interacts most strongly with thermal system because power dissipated by components generates heat, which must be dissipated by the spacecraft. Also, the amount of power available determines how much can be applied to heaters on components. In addition, power generation and storage devices, such as solar cells and batteries, often have stringent temperature requirements. For example, solar cell is inversely proportional to cell temperature, and solar panels can be sensitive to large variations in temperature that may deform the supporting structure, as occurred with the first generation of solar panels on Hubble.13

Among the five categories considered in this subsystem, orbital debris only interacts with the power generation system. If orbital debris were to impact and damage solar cells, for example, it would disrupt the amount of energy that can be collected.

Despite the sparse number of particles in LEO, the impacts of atmospheric molecules on a spacecraft can have varied effects. Atomic oxygen is a major component of the environment at altitudes between 200 and 800 km. The flux in this region is approximately 1019 atoms/m2-s.14 The flux of atmospheric oxygen causes measurable erosion on the surface of a spacecraft, sometimes as quickly a few days after deployment.14 Even in the absence of significant erosion, atomic oxygen changes the thermal properties of a spacecraft’s surface. Susceptibility to surface erosion by atomic oxygen varies largely by material. Even in the case of metals, which tend to have lower reactivities, microscopic effects must be considered due to the possible alteration of thermal properties.14

2.8 Intersystem Interactions

Figure 4 depicts the interactions between subsystems with arrows that vary in thickness proportionally with dependence. These arrows also indicate which subsystem in the interaction is dependent upon the other. If a spacecraft employs solar power generation techniques, the attitude determination and control system (ADCS) must account for the panel orientation, which can have an effect on moments of inertia. This orientation may be manipulated for optimum power generation, making the ADCS reliant on the power production system. The power system also depends on how the ADCS points the spacecraft and its solar panel. The power system will contribute to a portion of the mass and volume of the spacecraft, and as such must be taken into account by those designing the spacecraft structure. Some spacecraft use electricity to propel them, and so the propulsion system characteristics can be dependent on how much power the craft is able to generate. Communications systems require power to operate onboard a spacecraft, and so they too rely on the power system. If the power system is not fully automated, it must be given commands to execute specific functions, such as repositioning solar panels, discharging batteries, and when to draw power from the generator. These operations make the power system dependent on the communications system and ground station operators issuing and transmitting commands.

[pic]

Figure 4: Subsystem Interactions

The energy storage system is related to the power generation system because the energy stored is the excess energy produced. Also, the generation system should not produce more energy than the storage system can hold. The energy storage system can create heat when charging or discharging and therefore it affects the thermal system. Also, the energy storage system must be kept within a certain temperature range to operate optimally and this is one of the requirements of the thermal system. The discharging of the energy storage system produces the power for the entire satellite during eclipse. Therefore, the storage system’s capacity is defined by the power requirements of the satellite systems.

Other spacecraft components, in addition to solar panel structures, are sensitive to large temperature gradients. For example, cameras used in conjunction with ADCS often have strict pointing requirements that may be adversely affected by small structural deformations caused by thermal variations. Other structural components usually have larger acceptable temperature margins; however, the material used in the structure can affect the conductive and radiative heat transfer due to differing material properties such as thermal conductivity and absorptivity.

Components of the propulsion system often have tight temperature margins. Hydrazine, a commonly used propellant, has a low temperature limit based upon its freezing point and a high limit from its decomposition temperature. The communications and computer systems depend upon the thermal subsystem to keep important components within limits. The thermal subsystem is dependent upon the computer system to provide control for operational heaters and other thermal regulators, such as louvers. Finally, program management must provide the acceptable temperature margin data for all components, as well as the allowed cost and mass of the thermal subsystem.

Attitude dynamics and control systems strongly interact with orbital debris considerations. The design of the ADCS needs to include minor debris considerations in that its sensors and actuators need to be protected against impact by tiny particles in space.

The communications system is needed in small capacity to warn of large debris approaches to the satellite. This application would probably be used in conjunction with the ground systems, which can track large pieces of debris.

The structural system must withstand the impact of small particles that cannot be tracked. The orbit of the spacecraft and the amount of debris in that region should be considered when designing the structure.

The space environment causes several disturbing torques that must be taken into account by the attitude control system. The torques considered are due to gravity gradient, solar radiation, magnetic field, and aerodynamic disturbances. The values of these torques must be known to allow for accurate attitude control. These environmental torques are calculated using as follows.

The gravity gradient torque is found using:

[pic] (2.2-1)

where ( is the gravitational parameter, 3.986(1014 m3/s2, R the orbit radius, ( is the maximum angular deviation of the z-axis from the local vertical and Iz and Iy are the moments of inertia about the z- and y-axes.

The solar pressure torque is calculated by:

[pic] (2.2-2)

with [pic] (2.2-3)

where Fs is the solar constant, 1,367 W/m2, c is the speed of light, 3(108 m/s, [pic] is the angle of incidence of the Sun, cps ( cg is the distance between the center of solar pressure and the center of mass, and q is the coefficient of reflectivity.

The magnetic torque is found using:

[pic] (2.2-4)

where B is the Earth’s magnetic field in Tesla and D is the residual dipole of the spacecraft in A m2. B is calculated using Eq. (2.1-14).

Aerodynamic torque is calculated with the following equation:

[pic] (2.2-5)

with [pic] (2.2-6)

where F is the force in Newtons, ( is the atmospheric density in kg/m3, V the satellite velocity in m/s, A is the cross-sectional area of the spacecraft in the velocity direction in m2, Cd is the drag coefficient, and cpa ( cg is the distance between the center of aerodynamic pressure and the center of mass.

2.9 Summary and Overview

After performing preliminary analysis, the next step will be to identify the companies who manufacture the relevant components for this subsystem and to detail the specifications given for those components. The following chapter includes information about parts, manufacturers, examples of previous work in this area, and examples of calculations used to size the components.

Chapter 3: Subsystem Examples

This chapter introduces some examples of the systems described in previous chapters. These examples include the companies that manufacture each subsystem’s components as well as some discussion of the features of these devices. Also included are mathematical examples that show how parameters related to these subsystems are determined.

3.1 Power Generation Components

Many companies sell solar cells for use on space platforms today. Some of these companies include Tecstar27 and Emcore12. Tecstar offers multiple varieties of solar cells, such as Silicon BSR and BSF/R, single junction gallium arsenide (GaAs), dual junction cells, and their TEC3I Cascade triple junction solar cells. According to Tecstar, their Cascade multi-junction cells are qualified to operate in temperatures varying from -105( to +130(C and operate for 43,500 thermal cycles. Tecstar’s triple junction Cascade cells are rated at 23% efficiency. A single one square-meter panel of their GaInP2/GaAs/Ge cells provides 265 watts at spacecraft BOL, and 201 watts after being subjected to 1015 e/cm2 of radiation. These power calculations are based upon 4 cm ( 6 cm ( 140 (m cells covered with 4 mil antireflective coating CMX coverglass, with 478 cell assemblies per square-meter17. The mass of the cell stack for a one square-meter Cascade solar panel is 1.133 kg, and measures 1 m ( 1 m ( 433 (m in size. Tecstar has provided these cells to the following missions: Scarlet, SSTI/Clark, Mightysat, SSTI/Lewis, SMEX, and other classified satellites27.

Emcore is another solar cell manufacturer that makes triple and dual junction solar cells. Their triple junction cells are rated at 26% efficiency (BOL), and degrade less than 0.2% in efficiency after 2000 thermal cycles, each cycle from -180( to +95(C. Exposure to 1015 e/cm2 will result in a solar cell efficiency decrease to 22.6%. The mass of Emcore’s solar cells is listed as 86 mg/cm2, with one cell measuring 76.1 mm ( 37.2 mm ( 166 (m in size12.

Radioisotopic thermoelectric generators (RTGs) can also be a regenerative source of power aboard a spacecraft. These devices carry radioactive material that decays and produces heat, which in turn provides power to the spacecraft through the Seebeck effect. Fuels commonly used in a RTG include Polonium-210, Plutonium-238, Cesium-144, Strontium-90, and Curium-242.36 Typically, Plutonium-238 is used in RTGs due to its long half-life. RTGs weigh 55 kg, and contain 11 kg of plutonium dioxide fuel, pressed into 72 solid ceramic-like cylindrical 1- by 1-inch pellets.33 Each of the heat sources consists of 18 separate modules, each of which encases four of the Pu-238 pellets. These modules are designed to survive under a range of accidents, including launch vehicle explosion or fire, reentry into the atmosphere followed by land or water impact, and post-impact situations. RTG systems are very expensive, costing on the order of 3000 dollars per watt of thermal power, and some losses are present during the conversion to electrical power. Each RTG unit typically provides between 200 and 300 watts of power at the spacecraft beginning of life (BOL), but reduce output to between 150 and 200 watts after approximately 20 years.34 No special radiation shielding is necessary for these generators, however moderate neutron and gamma-ray fields exist external to the RTG, so they are typically isolated from the rest of the spacecraft. RTGs have been used in the past on such space missions as Galileo, the Ulysses Project, Pioneers 10 and 11, Voyagers 1 and 2, Cassini, and several of the manned Apollo missions.35 The Office of Special Nuclear Projects of the U.S. Department of Energy is responsible for the government RTG program.

3.2 Power Generation Example Calculations

One example of the application of solar power generation to a spacecraft is the design of a hypothetical craft: Spacecraft Z. This spacecraft is in a noon-midnight sun-synchronous orbit at an altitude of 400 km. It requires a steady 2000 watts of power, and an additional 200 watts during eclipse. The choice between solar power and thermoelectric power generation for this spacecraft must be based on the mission parameters. With the satellite in Earth orbit for its entire life, it is not necessary to use a RTG. In addition, each RTG can be quite costly and only provides between 200 and 300 watts at spacecraft BOL, so this mission would require many of them, becoming quite costly and massive. Solar panels are the optimum choice for this spacecraft’s needs. Using the process outlined in Section 2.2.1, we arrive at a solar panel area requirement of 24.14 m2 for gallium-arsenide cells.

3.3 Energy Storage Components

The method of storing energy in batteries has the longest heritage in space applications. The most commonly used space batteries are Nickel-Cadmium (Ni-Cd), Nickel-Hydride (Ni-MH), and Lithium-Ion (Li-ion). These batteries have different characteristics that can be catered to the specific needs of the design29. Space batteries are made by various manufacturers all over the world. To approximate the characteristics of battery storage devices the batteries from a single company are examined. This company is Haowei Industrial Corporation. The Haowei Co. manufactures Ni-Cd, Ni-MH, and Li-ion batteries.32

The Ni-Cd batteries from the Haowei Co. have many helpful features. They have high capacity density of more than 45 kW-hr/kg or more than 145 W-hr/dm3. They are rechargeable more than 500 full charge/discharge cycles. They have high, constant discharge voltages of 1.2 V per cell for 40-45 minutes in the discharge period at 1.0 C rate. They have a capability of high charge and discharge rates. They can be charged in 1-3 hours and can withstand a high rate of discharge up to 6-10 C. These batteries also have a wide operating temperature range of –20( C to 50( C. Nickel-Cadmium batteries have a memory effect that hinders the depth of charge. However, the Ni-Cd batteries from Haowei have the ability to remove the memory effect by a standard way of charging and discharging. Table 2 is a chart of the physical properties of the Ni-Cd batteries produced by Haowei.32

Table 1: Nickel Cadmium Cylindrical Rechargeable Battery32

|Cell Size |Model Number |Nominal Voltage |Nominal Capacity |Diameter (mm) |Height (mm) |Mass (g) |

| | |(V) |(mA-hr) | | | |

|AAA |HW150A AAAK |1.2 |350 |10 |43 |12 |

|AA |HW1000 AAK |1.2 |1000 |14 |49 |24 |

|A |HW1400 AK |1.2 |1400 |17 |49 |34 |

|C |HW2400 |1.2 |2400 |26 |50 |70 |

|D |HW5000 DKT |1.2 |5000 |32 |62 |150 |

|9V |HW100F 8K |8.4 |100 |L17.5 × W26.5 |48.5 |35 |

Haowei Co.’s Ni-MH batteries have advantages such as low self-discharge rates, rapid charge and discharge capabilities, and no memory effect. They have high capacity density of more than 55 kW-hr/kg or more than 190 kW-hr/dm3. They are rechargeable more than 500 full charge/discharge cycles. They have high, constant discharge voltages of 1.2 V per cell for 35-40 minutes in the discharge period at 1.0 C rate. They have a capability of high charge and discharge rates. They can be charged in 1-3 hours and can withstand a high rate of discharge up to 6-10 C. These batteries also have an operating temperature range of –20( C to 40( C. Table 3 is a chart of the physical properties of the Ni-MH batteries produced by Haowei.32

Table 3: Nickel Hydride Cylindrical Rechargeable Battery32

|Cell Size |Model Number |Nominal Voltage |Nominal Capacity (mA-hr) |Diameter (mm) |Height (mm) |Mass (g)|

| | |(V) | | | | |

|AAA |HW600A AAH |1.2 |600 |10 |43 |14 |

|AA |HW1500 AAH |1.2 |1500 |14 |49 |29 |

|A |HW2400 LAH |1.2 |2400 |17 |49 |45 |

|SC |HW2400 SCH |1.2 |2400 |23 |43 |57 |

Lithium-ion batteries are lightweight and have long cycle life and high voltage. They are capable of operating at more than 3 times the voltage of Ni-Cd and Ni-MH of 3.7 V. They have high capacity density of 250-350 kW-hr/kg or 450-550 kW-hr/dm3. They have a cycle life greater than 1000 full charge/discharge cycles. They also have no memory effects. These batteries also have an operating temperature range of –20( C to 60( C. Table 4 is a chart of the physical properties of the Li-ion batteries produced by Haowei.32

Table 4: Lithium-Ion Cylindrical Rechargeable Battery32

|Model Number |Nominal Voltage |Nominal Capacity (mA-hr) |Diameter (mm) |Height (mm) |Mass (g)|

| |(V) | | | | |

|HW C14500 |3.7 |580 |14 |50 |20 |

|HW C14650 |3.7 |750 |14 |65 |26 |

|HW C17500 |3.7 |800 |17 |50 |24 |

|HW C17670 |3.7 |1200 |17 |67 |35 |

|HW C18500 |3.7 |900 |18 |50 |26 |

|HW C18650 |3.7 |1550 |18 |65 |29 |

Flywheel energy storage is a new technology and therefore has not been widely implemented. There has been a lot of research in the area of flywheel energy storage. From the research and testing of this technology, flywheels have been proven to be a viable energy storage system.25

A flywheel is a ring or disk spinning on magnetic bearings in a vacuum. The limiting factor of flywheel energy storage is the tensile strength of the material in the ring/disk and its density. A material with high tensile strength and low density is able to store more energy. Therefore, researchers have looked at composites as possible material for the flywheels, namely Kevlar. Kevlar is a composite material that fractures into dust instead of fragmenting like a metal. This is a good characteristic for flywheels because failure protection is limiting factor in flywheel design. Kevlar has other good qualities such as a tensile strength of 4.8 GPa and a density of 1800 kg/m3.11

To prevent flywheel failure, flywheels are only spun to about 70% of their maximum rotational speed. This limits the amount of the energy that can but put into flywheels. However, flywheels usually return about 85% of the energy put into them.25 Typical flywheel energy storage systems have energy storage potential of approximately 181 W-hr/kg and have lifetimes of about 40 years. The prototypes produced by researchers produce 10 kW-hr of power. Based on these prototypes the cost of flywheel energy storage devices is estimated to be $500/kW-hr. However, new technology and methods developed in the future should reduce that figure by 20%.11

Thermal energy storage has been around for many years. However, because heat transfer is a slow process and because of inadequate insulation technology, thermal energy storage is a relatively inefficient process for space applications. Therefore, most of the research has gone into developing ground-based thermal energy storage systems.

Ground-based systems are more feasible because size and mass of the operating matter is not as important. This fact allows for large amounts working fluid or solid that store enough energy needed. In space, size and mass are critical in design of an energy storage system. Therefore, a thermal system would need to be lightweight, efficient, and cost effective. Material scientists are researching and developing new technology so that materials can store energy and processes can convert energy more efficiently. Ground-based systems have proven to be cost effective. They store and produce power at $0.002 /kW-hr. Therefore, thermal energy storage has the potential to be a viable solution if technology can provide better materials and processes.31

3.4 Thermal Subsystem Components

The most common components of any thermal subsystem are heaters, thermostats, thermal coatings, radiators, heat pipes, louvers, and blankets. This section gives a breakdown of these components, some of their important characteristics, as well as who manufactures them.

Heaters and thermostats are commonly implemented thermal system components. They are effective, relatively cheap, easy to manufacture, and small in mass. Heaters can range in size and power and are generally used to keep small spacecraft components within specified temperature ranges. Thermostats are used in conjunction with heaters to provide local temperature control for components. Both heaters and thermostats have been extensively used to provide thermal control on a wide assortment of space missions. Rosemount, Inc. manufactures space-rated thermostats and temperature sensors.24, 30

Thermal coatings are another widely implemented subsystem component. This is de partially to their versatility. Thermal coatings can vary greatly in type. Most are simply paints that can be applied to the surface of a spacecraft structure. Others are adhesives that have been treated on one side to have specific radiation characteristics. Thermal coatings are the least expensive of all thermal control options. One problem with thermal coatings is their limited lifetime. Most coatings will degrade over the span of a mission due to exposure to radiation and orbital debris. However, this degradation has been measured on previous missions and can be accounted for.13 Another problem with coatings is that many of them tend to outgas once placed in orbit. One must be careful to account for this when selecting coatings for the thermal system. The US company 3M manufactures many coatings and adhesives, some of which are space-rated and which have flown on missions.1

Radiators, heat pipes and louvers are examples of more expensive and complex devices used for thermal control. Radiators can vary greatly in size and dissipation capabilities from small panels on spacecraft instruments to the large radiator arrays on the International Space Station. Radiators and heat pipes have been extensively used on satellites and are thus a proven technology. Louvers are relatively new technology and have been tested recently on space shuttle missions. Swales Aerospace, a Maryland-based company, is a leading manufacturer of radiators, heat pipes and louvers.26

The use of mutli-layer insulation (MLI) can also provide adequate thermal control. This insulation is used to keep spacecraft components warm by preventing or damping radiation heat exchange to space. While simple in concept, MLI is somewhat difficult to manufacture because each piece must be hand crafted and fit checked to the component for which it is made. In addition, the layering material (usually silver coated Teflon) can be expensive. Aerospace Fabrication & Materials, LLC is a domestic company that specializes in the design, fabrication and installation of MLI.2

3.5 Thermal Subsystem Calculations

A spacecraft is to be placed into a 400 km circular Earth orbit. The spacecraft is rectangular with dimensions 1(1(4 m. A battery box is mounted to the nadir plate. The battery box has a footprint of 0.25 m2. An electronics box is mounted to the top plate opposite the battery box. The electronics box has a footprint of 0.25 m2 and dissipates 50 W of heat. The components of interest on this spacecraft are made of Aluminum alloy 6061-T6 and all walls are 0.1 m thick. The spacecraft is shown schematically in Figure 5.

[pic]

Figure 5: Schematic of Spacecraft Used for Example Calculations

The battery box has a maximum operational temperature of 40° C and a minimum of 0° C. For this example, we model the thermal characteristics of this satellite and design a thermal system that will keep the battery box within its temperature limits.

Since the spacecraft is in a circular orbit, we consider only two points in the orbit for our calculations: one where the satellite is in full sun, and one where it is in full eclipse. We ignore convective heat transfer since the spacecraft is far enough away from Earth’s atmosphere that frictional heating is negligible. In addition, we ignore internal radiation effects. Thus, only external radiation and internal conduction heat transfers need to be calculated. The necessary equations are 2.1-8 and 2.1-9.

First, we model the spacecraft as a series of thermal nodes. The nodal breakdown is shown in Figure 5. The battery and electronics boxes are one node each, and each wall of the satellite bus is one node. Next, we calculate the linear conduction couplings. The couplings are dependent upon the material properties of the structural material, which is in this case aluminum. 6061-T6 aluminum has a thermal conductivity (k) of 167.7 W/m-K.29 Conduction is also a property of the contact areas between the nodes of interest (A), as well as the conductive path length (Δx). The conduction couplings for each of the nodes are then calculated as shown in Table 5. Heat flows from node i to node j according to the magnitude of the conductor value given. Note that many of the conduction values are the same due to the symmetry of the model.

Table 5: Conduction Couplings Example

[pic]

Next, we can calculate the radiation couplings of the spacecraft. In order to do this, we require information about the surface properties of the external walls. The value of the radiation coupling will depend upon the emissivity (ε) of the coating chosen and the nodal surface area. For this example, we will consider that the entire surface is covered with silverized Teflon, which has an emissivity of 0.66.13 Because we are neglecting internal radiation effects, the only radiation couplings of interest here are those from the external walls to space (denoted here as Node 999). The radiation coupling calculations are shown in Table 6.

Table 6: Radiation Couplings Example

|Node i |Node j |Area (m2) |Radiation |Conduction |

| | | |(W/K4) |(W/K) |

|3 |999 |4.0 |0.25 |1.50(10-07 |

|4 |999 |4.0 |0.25 |1.50(10-07 |

|5 |999 |4.0 |0.5 |1.50(10-07 |

|6 |999 |4.0 |0.5 |1.50(10-07 |

|7 |999 |1.0 |0.5 |3.74(10-08 |

|8 |999 |1.0 |0.5 |3.74(10-08 |

Once the relevant radiation and conductive heat couplings are calculated, they can be used to determine the temperatures of the components of interest. A computer code such as SINDA or TRAYSYS can calculate both steady state and transient temperatures of modeled spacecraft components.13

3.6 Environment Components

3.6.1 Single Particle Events

The National Space Science Data Center (NSSDC) provides models for electron and proton flux. The electron and proton models are referred to as AE8 and AP8 respectively. The models also differentiate between solar min and solar max values. These models can be found at the NSSDC website.9

The Cosmic Ray Effects on Microelectronics (CREME) software package evaluates error rates due to cosmic ray bombardment on spacecraft microelectronics. The program is used to evaluate the differential and integral cosmic ray flux as a function of either particle energy or linear energy transfer (LET). The program can also estimate single event upset (SEU) rates and for microelectronics. CREME programs can be run from the CREME homepage.22 Inputs include the spacecraft’s orbit, the radiation environment being investigated, amount of shielding, and electronic device characteristics.

3.6.2 Electronic Devices

There are a number of parameters to be considered when selecting spacecraft microelectronics. These include: total dose effects, latchup threshold, upset threshold, SEU rates, and the component’s LET. When choosing parts, safety margins of 5 are recommended, margins of 2 to 5 are acceptable with additional testing, and margins of less than 2 should be avoided.14 In addition, devices with an LET threshold of greater than 100 MeV/mg/cm2 are recommended for minimizing SEUs. Other methods of safeguarding electronic systems include shielding, redundancy, and the use of recovery algorithms for protection against latchups.14

Space Electronics Incorporated offers a large variety of space-qualified electronics. Their products include: amplifiers, transceivers, microcontrollers and microprocessors, memory, switches, and multiplexers. Data sheets on their products can be found on their website.16 BAE Systems is another supplier of processors and electronics. Information on their products can be found on their website.7 is a web-based service listing a wide variety of products from suppliers around the world. Product specifications and links to manufacturers and suppliers are listed on their website.5

3.6.3 Debris and Thermal Protection

Multi-layer insulation (MLI) blankets are frequently used to protect spacecraft from a variety of environmental hazards. The company Spectrum Astro Inc. Makes MLI blankets for a wide range of applications including cryogenics, high temperature environments, micrometeoroid and orbital debris protection, and atomic oxygen erosion prevention.18 Austrian Aerospace produces several varieties of MLIs including ones made of Polyester and Mylar(, Polymide and Kapton(, titanium high temperature shielding, and high radiation shielding.6 Cyanate Ester is used in thermal management compounds applied to spacecraft hardware, heat sinks, and heat spreads. It has high thermal conductivity and low outgassing properties. Cyanate Ester also has a wide range of operating temperatures, -65( to 500( F. Bryte Technologies and JD Lincoln, Inc. produce Cyanate Ester.10, 16

3.6.4 Radiation Detection

Radiation exposure of a spacecraft can be measured by dosimeters. Amptek, Inc. manufactures the Compact Environmental Anomaly Sensor (CEASE). This device is capable of detecting total ionization radiation, radiation dose rating, surface charging, and single event effects. CEASE detects changes in the environment every 60 seconds. The device requires 1.5 to 1.7 W of power and its mass is between one and 1.3 kg.4

Traditionally, only the spacecraft’s structure (typically aluminum) and cargo were used to protect against radiation. This approach is no longer sufficient for sensitive electronics and other components. New candidate materials include borated polyethylene, which offers 6.3 times the protection of aluminum, and metal hydrides, which protect up to eight times as well as aluminum.3

3.7 Environment Design Calculations

Radiation dose rates vary with altitude. Figure 6 allows us to determine the required thickness of aluminum needed to protect spacecraft components from a specific radiation dose at a given altitude for low-altitude polar orbits.29 For instance, if shielding is needed up to 105 Rad/year at an altitude of 6000 km, then at least 0.5 g/cm2 of aluminum is needed. Similar graphs can be generated for other materials and orbits.

3.8 Summary and Overview

From this chapter, one should have an idea of how to choose and specify parts, and who makes these parts. While this information is not entirely comprehensive, it should serve as a starting point. An overall summary of the presented information and further recommendations are given in the next chapter.

Chapter 4: Conclusions

The information in the preceding chapters is not inclusive of all necessary issues, but should be used as a starting point for the design of a power, thermal, or space environment subsystem. This paper can also serve as a resource for relevant equations in those fields. A summary of what has been presented and the conclusions reached are provided in this chapter. Recommendations for further research are also given.

4.1 Summary and Conclusions

Power generation is a critical design aspect of any spacecraft. All systems onboard the craft must have power to function, making the power subsystem a vital spacecraft component. Choosing an appropriate power generation device is a delicate task aboard any spacecraft. Important things to consider when choosing a generation method are: spacecraft lifetime, where it will be traveling, and how much power is required. If the spacecraft is a deep space probe that will travel beyond Mars, RTGs become one of the optimum choices because they can continually generate power no matter where they’re located. If this same mission doesn’t have a prolonged life, stored energy devices such as batteries or fuel cells should be used due to their low cost. If the spacecraft’s mission keeps the satellite in orbit around the Earth, solar cells are the optimum choice because of their extensive heritage and low cost. Thus the choice of an ideal power generation method is highly dependent on the mission parameters.

The energy storage system takes energy generated by the spacecraft and stores it for use during periods of eclipse. Battery, flywheel, and thermal systems are examples of energy storage systems. Of these, batteries have been used the most in space systems. Batteries are inexpensive, low maintenance, and have high energy density. However the number of charge/discharge cycles and the DOD limit a battery’s lifetime. Batteries are also energy-rate dependent, and therefore constrain the amount of power brought into the system during charging. Flywheels are a new technology that have not yet been proven in space systems. However, they have longer lifetimes than batteries, can store energy at any rate necessary, and are efficient. Flywheels will be used extensively in the future for energy storage in space. Thermal systems for spacecraft use is a technology that is new and unproven. However, future thermal systems could be a cheap and viable technology for energy storage in space systems.

Spacecraft components must be maintained within their acceptable temperature margins in order to remain operational. Thus, the thermal subsystem must be designed with this objective in mind. The thermal system should be capable of providing adequate temperature control without drawing power from other systems. Ideally, the thermal system would be as passive as possible to minimize mass and power requirements.

The thermal subsystem is constrained by mass and temperature requirements as well as mission objectives such as orbit. For example, a spacecraft in a sun-synchronous orbit about the Earth will have significantly different thermal requirements than an interstellar mission to Jupiter. The thermal system is generally designed only after mass and orbit constraints have been defined. The thermal system also interacts strongly with the power subsystem. For instance, spacecraft batteries often have narrow temperature margins and the power available will constrain the amount of active thermal control possible.

The properties of the space environment need to be considered in designing a spacecraft because they can damage the structure, computers, and other systems. However, these properties are difficult to predict as they often vary with solar events. The effects of orbital debris, for example, are based largely on probabilities gathered from observations of previous missions.

Important devices to protect against environmental concerns include structural protection and detection devices. Both orbital debris and radiation can be absorbed by multi-layer insulation blankets. Careful consideration must be given to the selection of spacecraft microelectronics to safeguard against errors due to SEUs.

4.2 Recommendations

Future research efforts would be well spent on the enhancement solar cell technologies. The solar cell industry is constantly changing, and new advances are made quite often. Many new spacecraft concepts, such as catapult-style momentum exchange systems, require vast amounts of power for brief periods (on the order of 30 megawatts for 180 seconds). This presents a challenge for spacecraft power generation methods of today, as a system capable of these magnitudes of power generation would be extremely massive. It is important to explore the use of stored energy to supplement power generation. Energy stored during periods of excess power generation can help to lower the need for large generation systems by providing power for peak loads. For maximum efficiency, an optimum combination of power generation and energy storage must be used to provide the spacecraft’s power needs.

Future research into flywheels and thermal systems will help expand the possibilities of power systems. More flywheel research needs to be done in order to make flywheels more reliable and require less maintenance. Additional research into thermal systems must be done to make them more efficient and safe for spacecraft.

When designing a thermal subsystem for a spacecraft, several variables, such as component temperature, heat flux and conductive properties need to be considered. A method of modeling these variables was considered briefly in this report. However, before proceeding with more specific thermal designs, this method should be examined and refined.

A significant problem in designing spacecraft to withstand environmental effects is the lack of options. While there is research on a wide variety of materials, not many have been space-tested. Also, while many companies provide lists of environmental systems they manufacture, it is difficult to find cost and mass information on these systems. Such information is vital to incorporating the system into an actual spacecraft design.

In order to best safeguard a spacecraft from environmental effects, it is best to fully research the expected conditions that the spacecraft will be subjected to. This includes investigation of radiation models, orbital debris and micrometeoroid densities, and predicted solar event timelines. With this knowledge, a framework can be developed for protection from both collision and radiation damage.

References

1. 3M Worldwide. 3M Worldwide. Accessed October 7, 2001.



2. Aerospace Fabrication and Materials, LLC. AFM Home Page. Accessed October 6, 2001.



3. AIAA, Aerospace Sciences Meeting and Exhibit, 39th, Reno, NV, Jan. 8-11, 2001

4. Amptek. “CEASE” Compact Environmental Anomaly Sensor Space Radiation Alarm. Revised Aug. 28, 2001. Accessed Oct. 29, 2001.

cease.html

5. AstroWorks, Inc. The Space Industry Virtual Exhibit Hall. Updated October 29, 2000. Accessed October 5, 2001.



6. Austrian Aerospace. Multi-Layer Insulation Space Insulation. Accessed Oct. 22, 2001.

esa.int/est/prod/prod0674.htm

7. BAE Systems. Advanced Digital Solutions. Updated April 18, 2001. Accessed October 6, 2001.



8. Bedingfield, K.L., and R.D. Leach. Spacecraft System Failures and Anomalies Attributed to the Natural Space Environment, RP 1390. Ed. M.B. Alexander.

Alabama: NASA Marshall Space Flight Center. August 1996.

9. Bitlitza, Dieter, Dr. AE-8/AP-8 Radiation Belt Models. Updated January 18, 2001. National Space Science Data Center. Accessed October 6, 2001.

10. Bryte Technologies, Inc. Thermal Management Products. Last modified May 14, 2001. Accessed Oct. 23, 2001.



11. Dinkins, Jerry. Lee, Kerry. Mahserjian, Sam. Pinkoski, Bryon. Space Solar Power – Ground-Based Energy Storage. Texas Space Grant Consortium, Texas Tech University. December 10, 1997

12. EMCORE Corporation. Emcore Home. Accessed October 5, 2001.



13. Gilmore, David G., ed., Satellite Thermal Control Handbook, The Aerospace Corporation Press, El Segundo, CA, 1994.

14. Hastings, D. and H. Garrett, Spacecraft Environment Interactions, Cambridge University Press, UK, 1996.

15. Hubble: European Space Agency Information Centre. About Hubble. Revised April 27, 2000. Accessed October 7, 2001.



16. J.D. Lincoln, Inc. J.D. Lincoln, Inc. Accessed Oct. 23, 2001.



17. Jet Propulsion Laboratory. Spacecraft Power for NASA Missions. Accessed October 7, 2001.



18. Lepore, Albert. Thermal Products Center. Accessed Oct. 22, 2001.

19. Maxwell Technologies. Space Electronics. Accessed October 2, 2001.

20. NASA. Space Radiation Protection, NASA SP-8054. Langley Research Center, June 1970.

21. NASA. The Earth’s Ionosphere, NASA SP-8049. Goddard Space Flight Center, Greenbelt, Maryland, March 1971.

22. Naval Research Laboratory. Cosmic Ray Effects on Micro-Electronics. Updated September 20, 2001. Accessed October 5, 2001.

23. Reehorst, Sandra. Aerospace Flywheel Development. NASA Glenn Research Center. Accessed October 7, 2001.



24. Rosemount Inc. Rosemount Inc. Last Revised July 23, 2001. Accessed October 5, 2001.



25. Soeder, Flywheels Ready To Fly. Revised April 10, 2000. Accessed October 7, 2001. Cleveland

26. Swales Aerospace. Aerospace Products. Accessed October 5, 2001

27. Tecstar. Welcome to Tecstar! Accessed October 5, 2001.



28. Tribble, A. C. The Space Environment: Implications on Spacecraft Design, Princeton University Press, Princeton, NJ, 1995.

29. Wertz , James R. and Wiley J. Larson, ed., Space Mission Analysis and Design, Third Edition, Microcosm Press, El Segundo, CA, 1999.

30. Wilson, Andrew, ed. Jane’s Space Directory. Jane’s Information Group Limited. Coulsdon, Surrey, UK. 1995.

31. Wyman, Charles. Thermal Energy Storage for Solar Applications: An Overview. March 1979. Solar Energy Research Institute. pg. 25

32. Xin, Shenzhen. Haowei Industrial Co., Ltd. Revised 1999. Accessed October 7, 2001.



33. Are radioisotope thermoelectric generators safe? Accessed October 6, 2001.



34. Radioisotope Thermoelectric Generators (RTGs). From Issue 10, April 1984. Accessed October 7, 2001.



35. Spaceviews. The RTG Debate. Accessed October 6, 2001.

36. What are the fuels for radioisotope thermoelectric generators? Accessed October 5, 2001.



-----------------------

[pic]

Figure 6: Radiation Dose Rates as a Function of Altitude2

[pic]

................
................

In order to avoid copyright disputes, this page is only a partial summary.

Google Online Preview   Download